1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine airfoil with film cooling holes.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section that has a plurality of stages of stator vanes and rotor blades reacting to a high temperature gas flow passing through the turbine to convert the chemical energy from combustion into mechanical energy by rotating the turbine shaft. The efficiency of the turbine, and therefore of the engine, can be increased by increasing the hot gas flow that enters the turbine.
To allow for higher turbine entrance temperatures, the upper stage vanes and blades are made from exotic nickel alloys that can withstand very high temperatures and have complex internal cooling air passages to provide cooling to these airfoils. A thermal barrier coating (TBC) is also applied to the airfoil surfaces exposed to the hot gas flow in order to provide further protection from the heat. A TBC is typically made from a ceramic material. Also, the TBC is typically applied after the film cooling holes have been drilled into the airfoil surface to provide for the film cooling. These film cooling holes are limited to the diameter because of the drilling process.
Thicker TBC layers have been proposed to provide more protection to the airfoil substrate from the high temperature gas flow. As the TBC gets thicker, the thermal stresses developed in the TBC will tend to cause spalling.
It is therefore an object of the present invention to provide for an improved high temperature resistance coating applied to a turbine airfoil.
It is another object of the present invention to provide for a high temperature resistant coating with smaller diameter film cooling holes.
It is still another object of the present invention to provide for a process of forming small film cooling holes in a high temperature resistant coating on a turbine airfoil.